Rotor stall sensor system

ABSTRACT

A system for detecting onset of a stall in a rotor is disclosed, the system comprising a sensor spaced radially outwardly and apart from tips of a circumferential row of blades at a location on a static component that is between a first location and a second location wherein the first location is at a first distance of about 25% blade tip-chord length axially forward from the leading edge of a blade and the second location is at a second distance of about 25% blade tip-chord length axially aft from the trailing edge of a blade and wherein the sensor is capable of generating an input signal corresponding to a flow parameter at a location near the tip of a blade and indicative of the onset of a stall and a correlation processor that generates a stability correlation signal.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application is a Continuation-in-Part (CIP) patent application ofU.S. patent application Ser. No. 11/966,242, filed Dec. 28, 2007. Thecontents of that prior patent application are incorporated herein byreference in their entirety.

BACKGROUND OF THE INVENTION

This invention relates generally to gas turbine engines, and, morespecifically, to a system for detection of a stall in a compressionsystem therein.

In a turbofan aircraft gas turbine engine, air is pressurized in acompression system, comprising a fan module, a booster module and acompression module during operation. In large turbo fan engines, the airpassing through the fan module is mostly passed into a by-pass streamand used for generating the bulk of the thrust needed for propelling anaircraft in flight. The air channeled through the booster module andcompression module is mixed with fuel in a combustor and ignited,generating hot combustion gases which flow through turbine stages thatextract energy therefrom for powering the fan, booster and compressorrotors. The fan, booster and compressor modules have a series of rotorstages and stator stages. The fan and booster rotors are typicallydriven by a low pressure turbine and the compressor rotor is driven by ahigh pressure turbine. The fan and booster rotors are aerodynamicallycoupled to the compressor rotor although these normally operate atdifferent mechanical speeds.

Operability in a wide range of operating conditions is a fundamentalrequirement in the design of compression systems, such as fans, boostersand compressors. Modern developments in advanced aircrafts have requiredthe use of engines buried within the airframe, with air flowing into theengines through inlets that have unique geometries that cause severedistortions in the inlet airflow. Some of these engines may also have afixed area exhaust nozzle, which limits the operability of theseengines. Fundamental in the design of these compression systems isefficiency in compressing the air with sufficient stall margin over theentire flight envelope of operation from takeoff, cruise, and landing.However, compression efficiency and stall margin are normally inverselyrelated with increasing efficiency typically corresponding with adecrease in stall margin. The conflicting requirements of stall marginand efficiency are particularly demanding in high performance jetengines that operate under challenging operating conditions such assevere inlet distortions, fixed area nozzles and increased auxiliarypower extractions, while still requiring high a level of stabilitymargin throughout the flight envelope.

Stalls are commonly caused by flow breakdowns at the tip of the rotorblades of compression systems such as fans, compressors and boosters. Ingas turbine engine compression system rotors, there are tip clearancesbetween rotating blade tips and a stationary casing or shroud thatsurrounds the blade tips. During the engine operation, air leaks fromthe pressure side of a blade through the tip clearance toward thesuction side. These leakage flows may cause vortices to form at the tipregion of the blade. A tip vortex can grow and spread when there aresevere inlet distortions in the air flowing into compression system orwhen the engine is throttled and lead to a compressor stall and causesignificant operability problems and performance losses.

Accordingly, it would be desirable to have the ability to measure andcontrol dynamic processes such as flow instabilities in a compressionsystem. It would be desirable to have a system that can measure aparameter related to the onset of flow instabilities, such as thedynamic pressure, temperature, velocity and/or entropy near the bladetips, and process the measured data to predict the onset of stall in astage of a compression system, such as a fan or compressor. It wouldalso be desirable to have a system to mitigate compression system stallsbased on the measurement system output, for certain flight maneuvers atcritical points in the flight envelope, allowing the maneuvers to becompleted without stall or surge.

BRIEF DESCRIPTION OF THE INVENTION

The above-mentioned need or needs may be met by exemplary embodimentswhich provide a system for detecting onset of a stall in a rotor, thesystem comprising a sensor spaced radially outwardly and apart from tipsof a circumferential row of blades at a location on a static componentthat is between a first location and a second location wherein the firstlocation is at a first distance of about 25% blade tip-chord lengthaxially forward from the leading edge of a blade and the second locationis at a second distance of about 25% blade tip-chord length axially aftfrom the trailing edge of a blade and wherein the sensor is capable ofgenerating an input signal corresponding to a flow parameter at alocation near the tip of a blade and indicative of the onset of a stalland a correlation processor that generates a stability correlationsignal.

In another embodiment, a system for detecting onset of a stall comprisesa pressure sensor capable of generating a signal corresponding to thepressure at a location near the blade tip.

In another embodiment, a system for detecting onset of a stall comprisesa temperature sensor capable of generating a signal corresponding to thetemperature at a location near the blade tip.

In another embodiment, a system for detecting onset of a stall comprisesa velocity sensor capable of generating a signal corresponding to thevelocity at a location near the blade tip.

In another embodiment, a system for detecting onset of a stall comprisesan entropy sensor capable of generating a signal corresponding to theentropy at a location near the blade tip.

BRIEF DESCRIPTION OF THE DRAWINGS

The subject matter which is regarded as the invention is particularlypointed out and distinctly claimed in the concluding part of thespecification. The invention, however, may be best understood byreference to the following description taken in conjunction with theaccompanying drawing figures in which:

FIG. 1 is a schematic cross-sectional view of a gas turbine engine withan exemplary embodiment of the present invention.

FIG. 2 is an enlarged cross-sectional view of a portion of the fansection of the gas turbine engine shown in FIG. 1.

FIG. 3 is an exemplary operating map of a compression system in the gasturbine engine shown in FIG. 1.

FIG. 4 a shows the formation of a region with blade tip vortex in a fanstage.

FIG. 4 b shows the spread of the blade tip vortex shown in FIG. 4 a.

FIG. 4 c shows the vortex flow at blade tip region during a stall.

FIG. 5 is a schematic cross-sectional view of the tip region of a fanwith an exemplary embodiment of a stall detection system.

FIG. 6 is a schematic sketch of an exemplary arrangement of multiplesensors for a stall detection system.

FIG. 7 is a schematic sketch of exemplary locations of sensors in arotor stall sensor system.

FIG. 8 is an exemplary time history of pressure from an unsteady CFDsimulation of a compression system approaching stall.

FIG. 9 is an exemplary time history of temperature from an unsteady CFDsimulation of a compression system approaching stall.

FIG. 10 is an exemplary time history of velocity from an unsteady CFDsimulation of a compression system approaching stall.

FIG. 11 is an exemplary time history of entropy from an unsteady CFDsimulation of a compression system approaching stall.

DETAILED DESCRIPTION OF THE INVENTION

Referring to the drawings wherein identical reference numerals denotethe same elements throughout the various views, FIG. 1 shows anexemplary turbofan gas turbine engine 10 incorporating an exemplaryembodiment of the present invention. It comprises an engine centerlineaxis 8, fan section 12 which receives ambient air 14, a high pressurecompressor (HPC) 18, a combustor 20 which mixes fuel with the airpressurized by the HPC 18 for generating combustion gases or gas flowwhich flows downstream through a high pressure turbine (HPT) 22, and alow pressure turbine (LPT) 24 from which the combustion gases aredischarged from the engine 10. Many engines have a booster or lowpressure compressor (not shown in FIG. 1) mounted between the fansection and the HPC. A portion of the air passing through the fansection 12 is bypassed around the high pressure compressor 18 through abypass duct 21 having an entrance or splitter 23 between the fan section12 and the high pressure compressor 18. The HPT 22 is joined to the HPC18 to substantially form a high pressure rotor 29. A low pressure shaft28 joins the LPT 24 to the fan section 12 and the booster if one isused. The second or low pressure shaft 28 is rotatably disposedco-axially with and radially inwardly of the first or high pressurerotor. In the exemplary embodiments of the present invention shown inFIGS. 1 and 2, the fan section 12 has a multi-stage fan rotor, as inmany gas turbine engines, illustrated by first, second, and third fanrotor stages 12 a, 12 b, and 12 c respectively.

The fan section 12 that pressurizes the air flowing through it isaxisymmetrical about the longitudinal centerline axis 8. The fan section12 includes a plurality of inlet guide vanes (IGV) 30 and a plurality ofstator vanes 31 arranged in a circumferential direction around thelongitudinal centerline axis 8. The multiple fan rotor stages 12 of thefan section 12 have corresponding fan rotor blades 40 a, 40 b, 40 cextending radially outwardly from corresponding rotor hubs 39 a, 39 b,39 c in the form of separate disks, or integral blisks, or annular drumsin any conventional manner.

Cooperating with a fan rotor stage 12 a, 12 b, 12 c is a correspondingstator stage comprising a plurality of circumferentially spaced apartstator vanes 31 a, 31 b, 31 c. The arrangement of stator vanes and rotorblades is shown in FIG. 2. The rotor blades 40 and stator vanes 31define airfoils having corresponding aerodynamic profiles or contoursfor pressurizing the airflow successively in axial stages. Each fanrotor blade 40 comprises an airfoil 34 extending radially outward from ablade root 45 to a blade tip 46, a pressure side 43, a suction side 44,a leading edge 41 and a trailing edge 42. The airfoil 34 extends in thechordwise direction between the leading edge 41 and the trailing edge42. A chord C of the airfoil 34 is the length between the leading 41 andtrailing edge 42 at each radial cross section of the blade. The pressureside 43 of the airfoil 34 faces in the general direction of rotation ofthe fan rotors and the suction side 44 is on the other side of theairfoil. The front stage rotor blades 40 rotate within an annular casing50 that surrounds the rotor blade tips. The aft stage rotor bladestypically rotate within an annular passage formed by shroud segments 51that are circumferentially arranged around the blade tips 46. Inoperation, pressure of the air is increased as the air decelerates anddiffuses through the stator and rotor airfoils.

Operating map of an exemplary compression system, such as the fansection 12 in the exemplary gas turbine engine 10 is shown in FIG. 3,with inlet corrected flow rate along the X-axis and the pressure ratioon the Y-axis. Operating lines 114, 116 and the stall line 112 areshown, along with constant speed lines 122, 124. Line 124 represents alower speed line and line 122 represents a higher speed line. As thecompression system is throttled at a constant speed, such as constantspeed line 124, the inlet corrected flow rate decreases while thepressure ratio increases, and the compression system operation movescloser to the stall line 112. Each operating condition has acorresponding compressor efficiency, conventionally defined as the ratioof ideal (isentropic) compressor work input to actual work inputrequired to achieve a given pressure ratio. The compressor efficiency ofeach operating condition is plotted on the operating map in the form ofcontours of constant efficiency, such as items 118, 120 shown in FIG. 3.The performance map has a region of peak efficiency, depicted in FIG. 3as the smallest contour 120, and it is desirable to operate thecompression systems in the region of peak efficiency as much aspossible. Flow distortions in the inlet air flow 14 which enters the fansection 12 tend to cause flow instabilities as the air is compressed bythe fan blades (and compression system blades) and the stall line 112will tend to drop lower. As explained further below herein, theexemplary embodiments of the present invention provide a system fordetecting the flow instabilities in the fan section 12, such as fromflow distortions, and processing the information from the fan section topredict an impending stall in a fan rotor. The embodiments of thepresent invention shown herein enable other systems in the engine whichcan respond as necessary to manage the stall margin of fan rotors andother compression systems.

Stalls in fan rotors due to inlet flow distortions, and stalls in othercompression systems that are throttled, are known to be caused by abreakdown of flow in the tip region 52 of rotors, such as, for example,the fan rotors 12 a, 12 b, 12 c shown in FIG. 2. This tip flow breakdownis associated with tip leakage vortex schematically shown in FIGS. 4 a,4 b and 4 c as contour plots of regions having a negative axialvelocity, based from computational fluid dynamic analyses. Tip leakagevortex 200 initiates primarily at the rotor blade tip 46 near theleading edge 41. In the region of this vortex 200, there exists flowthat has negative axial velocity, that is, the flow in this region iscounter to the main body of flow and is highly undesirable. Unlessinterrupted, the tip vortex 200 propagates axially aft and tangentiallyfrom the blade suction surface 44 to the adjacent blade pressure surface43 as shown in FIG. 4 b. Once it reaches the pressure surface 43, theflow tends to collect in a region of blockage at the tip between theblades as shown in FIG. 4 c and causes high loss. As the inlet flowdistortions become severe, or as a compression system is throttled, theblockage becomes increasingly larger within the flow passage between theadjacent blades and eventually becomes so large as to drop the rotorpressure ratio below its design level, and causes the fan rotor tostall. Near stall, the behavior of the blade passage flow fieldstructure, specifically the blade tip clearance vortex trajectory, isperpendicular to the axial direction wherein the tip clearance vortex200 spans the leading edges 41 of adjacent blades 40, as shown in FIG. 4c. The vortex 200 starts from the leading edge 41 on the suction surface44 of the blade 40 and moves towards the leading edge 41 on the pressureside of the adjacent blade 40 as shown in FIG. 4 c.

The ability to control a dynamic process, such as a flow instability ina compression system, requires a measurement of a characteristic of theprocess. A continuous measurement or samples of sufficient number ofdiscrete measurements. In order to mitigate compression system stallsfor certain flight maneuvers at critical points in the flight envelopewhere the stability margin is small or negative, a flow parameter in theengine is first measured that can be used directly or, with someadditional processing, to predict the onset of stall of a stage of acompression system, such as, for example, a multistage fan shown in FIG.2.

FIG. 2 shows an exemplary embodiment of a system 500 for detecting theonset of an aerodynamic instability, such as a stall or surge, in acompression stage in a gas turbine engine 10. In the exemplaryembodiment shown in FIG. 2, a fan section 12 shown, comprising a threestage first rotor, 12 a, 12 b and 12 c. The embodiments of the presentinvention can also be used in a single stage fan, or in othercompression system in a gas turbine engine, such as a high pressurecompressor 18 or a low pressure compressor or a booster. In theexemplary embodiments shown herein, a sensor 502 is used to measure alocal flow property near the tip region 52 of the compression systemrotor blade tips 46 during engine operation. Although a single sensor502 can be used for the flow parameter measurements, use of at least twosensors 502 is preferred, because some sensors may become inoperableduring extended periods of engine operations. In an exemplary embodimentshown in FIG. 2, multiple sensors 502 are used around the tips of allthree fan rotor stages 12 a, 12 b, and 12 c.

In the exemplary embodiment shown in FIG. 5, the sensor 502 is locatedon a casing 50 that is spaced radially outwardly and apart from the fanblade tips 46. Alternatively, the sensor 502 may be located on a shroudsegment 51 that is located radially outwards from the blade tips. Thecasing 50, or a plurality of shrouds 51, surrounds the tips of a row ofblades 47. The sensors 502 are arranged circumferentially on the casing50 or the shrouds 51, as shown in FIG. 6. In an exemplary embodiment ofusing multiple sensors on a rotor stage, the sensors 502 are arranged insubstantially diametrically opposite locations in the casing or shroud.

During engine operation, there is an effective clearance 48 between therotor blade tip and the casing 50 or the shroud 51 (see FIG. 5). Thesensor 502 is capable of generating an input signal 504 in real timecorresponding to a flow parameter, such as, for example, the dynamicpressure, temperature, velocity and/or entropy in the blade tip region52 near the blade tip 46. A suitable high response transducer, having aresponse capability higher than the blade passing frequency is used.Typically these transducers have a response capability higher than 1000Hz. In the preferred embodiment shown herein the sensors 502 used weredynamic pressure sensors 202 made by Kulite Semiconductor Products. Inan alternative embodiment, the sensor 502 is any commercially availabletemperature sensor 204 having suitable dynamic capabilities. In anotheralternative embodiment, the sensor 502 is any commercially availablevelocity sensor 206 having suitable dynamic capabilities. In anotheralternative embodiment, the sensor 502 is any commercially availableentropy sensor 208 having suitable dynamic capabilities. It ispreferable to use a high frequency sampling of the flow propertiesmeasurements, such as for example, between ten and twenty times theblade passing frequency.

The flow parameter measurement from the sensor 502 generates a signalthat is used as an input signal 504 by a correlation processor 510. Thecorrelation processor 510 may also optionally receive as input a signal506 corresponding to the rotational speed of the compression systemrotor, such as the fan rotor 12 a, 12 b, 12 c, as shown in FIGS. 1, 2and 5. The signal 506 indicative of the rotational speed of the rotor(referred to herein as rotational speed signal 506 or as rotor speedsignal 506) may be generated by any known methods, such as usingrotational speed sensors, blade proximity sensors, or other knowndevices and methods. The rotor speed signal, when used, can provide onemethod of determining the blade passing period and/or frequency.However, such a measurement of rotor speed is not necessarily arequirement. Blade passing period/frequency can also be determined inreal-time, using known methods, from the signal from an unsteadypressure, temperature, velocity or entropy sensor near the blade tip. Inthe exemplary embodiments shown herein, the compression system rotorspeed signal 506 is supplied by a conventional engine control system 74,that is used in gas turbine engines. Alternatively, the compressionsystem rotor speed signal 506 may be supplied by a digital electroniccontrol system or a Full Authority Digital Electronic Control (FADEC)system used an aircraft engine.

The correlation processor 510 receives the input signal 504 from thesensor 502 and the rotor speed signal 506 from the control system 74 andgenerates a stability correlation signal 512 in real time usingconventional numerical methods. Auto correlation methods available inthe published literature may be used. In the exemplary embodiments shownherein, the correlation processor 510 algorithm uses the existing speedsignal from the engine control for cycle synchronization. Thecorrelation measure is computed for individual sensors over rotor bladetips. The auto-correlation system in the exemplary embodiments describedherein sampled a signal from a sensor 502 at a frequency of 200 KHz. Arelatively high value of sampling frequency in the range of about200-400 KHz ensures that the data is sampled at a rate at least tentimes the fan blade 40 passage frequency. A window of seventy twosamples was used to calculate the auto-correlation showing a value ofnear unity along the operating line 116 and dropping towards zero whenthe operation approached the stall/surge line 112 (see FIG. 3). For aparticular compression system rotor stage, such as, for example, the fanstage 12 a, 12 b, 12 c, when the stability margin approaches zero, theparticular compression stage rotor is on the verge of stall and thecorrelation measure is at a minimum. In systems designed to avoid astall or surge in a compression system, when the correlation measuredrops below a selected and pre-set threshold level, a stabilitymanagement system receives the stability correlation signal 512 andsends an electrical signal to the engine control system, such as forexample a FADEC system, which in turn can take corrective action usingthe available control devices to move the engine away from surge. Themethods used by the correlation processor 510 for gauging theaerodynamic stability level in the exemplary embodiment shown herein isdescribed in the paper, “Development and Demonstration of a StabilityManagement System for Gas Turbine Engines”, Proceedings of GT2006 ASMETurbo Expo 2006, GT2006-90324.

FIG. 5 shows schematically an exemplary embodiment of the presentinvention using a sensor 502 located in a casing 50 near the blade tipmid-chord of a blade 40. The sensor is located in the casing 50 suchthat it can measure a flow property of the air in the clearance 48between a rotor blade tip 46 and the inner surface 53 of the casing 50.In one exemplary embodiment, the sensor 502 is located in an annulargroove 54 in the casing 50. In other exemplary embodiments, it ispossible to have multiple annular grooves 54 in the casing 50, such asfor example, to provide for tip flow modifications for stability. Ifmultiple grooves are present, the sensor 502 is located within some ofthese grooves, using the same principles and examples disclosed herein.Although the sensor is shown in FIG. 5 as located in a casing 50, inother embodiments, the sensor 502 may be located in a shroud 51 that islocated radially outwards and apart from the blade tip 46. Also, thesensor 502 may be located in a casing 50 (or shroud 51) near the leadingedge 41 tip or the trailing edge 42 tip of the blade 40.

FIG. 6 shows schematically an exemplary embodiment of the presentinvention using a plurality of sensors 502 in a compression system, suchas a fan stage, shown in FIG. 2. The plurality of sensors 502 arearranged in the casing 50 (or shroud 51) in a circumferential direction,such that pairs of sensors 502 are located substantially diametricallyopposite. The correlations processor 510 receives input signals 504 fromthese pairs of sensors and processes signals from the pairs together.The differences in the measured data from the diametrically oppositesensors in a pair can be particularly useful in developing stabilitycorrelation signal 512 to detect the on set of a fan stall due to engineinlet flow distortions. A single sensor has been demonstrated to besufficient in some applications.

FIG. 7 shows the axial location of the sensors 502, such as the pressuresensor 202, temperature sensor 204, velocity sensor 206, entropy sensor208 or a plasma sensor 60 with respect to the compression system rotorblade leading edge 41 and trailing edge 42. In a particular application,it is possible to have any one or more types of these flow propertysensors. It is not necessary to have all these sensors in particularapplication and a suitable combination to obtain optimum results may beused. In FIG. 7, the rotor blade tip chord 49 is shown labeled “C”. Thetip chord C of the airfoil 34 is the axial length between the leading 41and trailing edge 42 at the tip of the blade. In the present invention,the sensor 502 (such as the pressure sensor 202, temperature sensor 204,velocity sensor 206, entropy sensor 208 and plasma sensor 60) is locatedradially outwardly and apart from tips 46 of a circumferential row ofblades 47 at a location on a static component 50 (such as a casing or ashroud) that is between a first location 57 and a second location 58.See FIG. 7. The first location 57 is at a first distance 157 (labeled as“A”) of about 25% blade tip-chord length 49 (“C”) axially forward fromthe leading edge 41 of a blade 47. The second location 58 is at a seconddistance 158 (labeled as “B”) of about 25% blade tip-chord length 49(“C”) axially aft from the trailing edge 42 of a blade 47. Thus thesensor 502 may be located at a suitable axial location in the region 159(labeled “D” in FIG. 7). In a preferred embodiment, the sensor islocated at an axial location corresponding to the mid-chord of the rotorblade tip.

FIGS. 8-11 show time history of the flow properties, pressure,temperature, velocity and entropy from an unsteady computational fluiddynamic (CFD) simulation in the rotor's relative frame of reference asthe compression system approaches a stall condition. Testing experiencehas demonstrated that unsteady pressure measurements can be successfullyused for autocorrelation calculations to predict an impending stallcondition. As discussed previously herein, a lack of correlation betweensuccessive measurements in a rotor is observed when a stall isapproaching. As evident in the unsteady CFD simulation shown in FIG. 8,the three local dips (items 302) below the zero non-dimensional pressureare typical of features that result in low autocorrelation when observedin the absolute frame of reference. Known autocorrelation algorithms canbe used on the pressure measurements from the pressure sensor 202.

FIG. 9 shows the time history of the temperature at the location of atemperature sensor 204 in an alternative embodiment of the presentinvention from an unsteady computational fluid dynamic (CFD) simulationin the rotor's relative frame of reference as the compression systemapproaches a stall condition. In this alternative embodiment, unsteadytemperature measurements using the temperature sensor 204 can be usedfor autocorrelation calculations to predict an impending stallcondition. As discussed previously herein, a lack of correlation betweensuccessive measurements in a rotor is observed when a stall isapproaching. As evident in the unsteady CFD simulation shown in FIG. 9,the three local peaks (items 304) above the zero non-dimensionaltemperature are typical of features that result in low autocorrelationwhen observed in the absolute frame of reference. Known autocorrelationalgorithms can be used on the temperature measurements from thetemperature sensor 204.

FIG. 10 shows the time history of the air velocity at the location of avelocity sensor 206 in an alternative embodiment of the presentinvention from an unsteady computational fluid dynamic (CFD) simulationin the rotor's relative frame of reference as the compression systemapproaches a stall condition. In this alternative embodiment, unsteadyvelocity measurements using the velocity sensor 206 can be used forautocorrelation calculations to predict an impending stall condition. Asdiscussed previously herein, a lack of correlation between successivemeasurements in a rotor is observed when a stall is approaching. Asevident in the unsteady CFD simulation shown in FIG. 10, the three localdips (items 306) below the zero non-dimensional velocity are typical offeatures that result in low autocorrelation when observed in theabsolute frame of reference. Known autocorrelation algorithms can beused on the velocity measurements from the velocity sensor 206.

FIG. 11 shows the time history of the entropy at the location of anentropy sensor 208 in an alternative embodiment of the present inventionfrom an unsteady computational fluid dynamic (CFD) simulation in therotor's relative frame of reference as the compression system approachesa stall condition. In this alternative embodiment, unsteady entropymeasurements using the entropy sensor 208 can be used forautocorrelation calculations to predict an impending stall condition. Asdiscussed previously herein, a lack of correlation between successivemeasurements in a rotor is observed when a stall is approaching. Asevident in the unsteady CFD simulation shown in FIG. 11, the three localpeaks (items 308) above the zero non-dimensional entropy are typical offeatures that result in low autocorrelation when observed in theabsolute frame of reference. Known autocorrelation algorithms can beused on the entropy measurements from the entropy sensor 308.

This written description uses examples to disclose the invention,including the best mode, and also to enable any person skilled in theart to make and use the invention. The patentable scope of the inventionis defined by the claims, and may include other examples that occur tothose skilled in the art. Such other examples are intended to be withinthe scope of the claims if they have structural elements that do notdiffer from the literal language of the claims, or if they includeequivalent structural elements with insubstantial differences from theliteral languages of the claims.

1. A system for detecting onset of a stall in a rotor, the systemcomprising: a sensor spaced radially outwardly and apart from tips of acircumferential row of blades at a location on a static component thatis between a first location and a second location wherein the firstlocation is at a first distance of about 25% blade tip-chord lengthaxially forward from the leading edge of a blade and the second locationis at a second distance of about 25% blade tip-chord length axially aftfrom the trailing edge of a blade and wherein the sensor is capable ofgenerating an input signal corresponding to a flow parameter at alocation near the tip of a blade and indicative of the onset of a stall;a control system capable of generating a rotor speed signal; and acorrelation processor capable of receiving the input signal and therotor speed signal wherein the correlation processor generates astability correlation signal.
 2. A system according to claim 1 furthercomprising: a plurality of sensors arranged on the static componentspaced radially outwardly and apart from tips of the row of blades.
 3. Asystem according to claim 1 wherein the sensor is a pressure sensorcapable of generating a signal corresponding to the pressure at alocation near the blade tip.
 4. A system according to claim 1 whereinthe sensor is a temperature sensor capable of generating a signalcorresponding to the temperature at a location near the blade tip.
 5. Asystem according to claim 4 wherein the sensor is located at a locationon the static component corresponding to the mid-chord of a blade.
 6. Asystem according to claim 1 wherein the sensor is a velocity sensorcapable of generating a signal corresponding to the flow velocity at alocation near the blade tip.
 7. A system according to claim 1 whereinthe sensor is capable of generating a signal corresponding to theentropy at a location near the blade tip.
 8. A system according to claim1 further comprising: a plurality of sensors arranged circumferentiallyon the static component around an axis of rotation of the rotor andspaced radially outwardly and apart from tips of the row of blades.
 9. Asystem according to claim 1 wherein the static component is a casing.10. A system according to claim 1 wherein the static component is ashroud.
 11. A system according to claim 1 wherein the rotor comprises aplurality of fan rotors.
 12. A system according to claim 1 wherein therotor is a compressor rotor.
 13. A system according to claim 1 whereinthe rotor is a booster rotor.
 14. A system for detecting onset of astall in a compressor rotor comprising: a sensor spaced radiallyoutwardly and apart from tips of a circumferential row of compressorblades at a location on a static component that is between a firstlocation and a second location wherein the first location is at a firstdistance of about 25% blade tip-chord length axially forward from theleading edge of a blade and the second location is at a second distanceof about 25% blade tip-chord length axially aft from the trailing edgeof a blade and wherein the sensor is capable of generating an inputsignal corresponding to a flow parameter at a location near the tip of ablade and indicative of the onset of a stall in the compressor; and acorrelation processor capable of receiving the input signal and a rotorspeed signal wherein the correlation processor generates a stabilitycorrelation signal.
 15. A system according to claim 14 furthercomprising a plurality of compressor rotors wherein a plurality sensorsare located on the static component surrounding tips of compressorblades of at least two compressor rotors.
 16. A system according toclaim 14 further comprising a plurality of sensors arrangedcircumferentially on the static component around an axis of rotation ofthe compressor rotor and spaced radially outwardly and apart from tipsof the row of compressor blades.
 17. A system according to claim 14wherein the sensor is a pressure sensor capable of generating a signalcorresponding to the pressure at a location near the compressor bladetip.
 18. A system according to claim 14 wherein the sensor is atemperature sensor capable of generating a signal corresponding to thetemperature at a location near the compressor blade tip.
 19. A systemaccording to claim 14 wherein the sensor is a velocity sensor capable ofgenerating a signal corresponding to the flow velocity at a locationnear the compressor blade tip.
 20. A system according to claim 14wherein the sensor is capable of generating a signal corresponding tothe entropy at a location near the compressor blade tip.